Systems and methods to prevent cracking of exterior paint along structural joints of painted aerospace components

ABSTRACT

Anti-paint cracking systems are provided as part of a painted aerospace component (e.g., a fuselage, an engine nacelle, a wing, an empennage, a fairing and the like) having a pair of adjacent structural panels defining therebetween a joint space having a lengthwise extent and a finished paint topcoat covering an exterior surface of the adjacent structural panels and the joint space therebetween. The anti-paint cracking system will thus be associated with the joint space along the lengthwise extent thereof and positioned below the finished paint topcoat. The anti-paint cracking system will include a joint sealant positioned in the joint space and an adhesive tape or a fibrous tape formed of a fiber composite material applied over the joint space and the joint sealant positioned therein.

FIELD

The embodiments described herein relate generally to aerospacecomponents having a finished exterior paint layer covering thestructural joints thereof (e.g., joints between adjacent aircraftfuselage skin panels), and especially relate to systems and methodswhich prevent cracking of such finished paint layer at the structuraljoints of the aerospace component.

BACKGROUND

Cracks in the finished paint on sealed structural joints associated withaerospace components (e.g., joints between adjacent panels associatedwith a fuselage, engine nacelles, wing skin panels, empennage panels,fairings and the like) has been an issue in the industry which adverselyimpacts the visual aspects of the finished paint scheme. The issue isparticularly acute for the fuselage paint on executive jets where a highquality and visually pleasing paint finish is an important customerrequirement. However, cracks in the finished paint along structuraljoints of the fuselage associated with commercial aviation aircraft isalso a very common problem.

Cracks in the finished fuselage paint occur due to the difference ofmechanical properties of the materials used at the structural joints(e.g., differences in mechanical properties of the sealant, the metalpanel (typically aluminum), and finish paint). These mechanical propertydifferences in turn create localized strain on the finished paint layerwhen the structure experiences a load, e.g., due to forces caused byinternal fuselage pressurization forces and external aerodynamic forces.The consequence of such forces is that the paint visibly cracks alongthe lengthwise extent of the joint.

It has previously been proposed in this art to create a gap in thesealant between the structural panels before applying the final paintcoating. However, this proposed solution has limitations due to thevisual aspect of the gap and lack of aerodynamic smoothness. As such,this prior proposal has essentially been limited for use mainly onfairings joining the wings to the fuselage and maintenance access doors.

What has been needed in the art, therefore is a system by which thefinished paint covering structural joints of an aerospace component(e.g., an aircraft fuselage) could be protected against cracking. It istowards fulfilling such a need that the embodiments disclosed herein aredirected.

SUMMARY

In general, the anti-paint cracking systems will be provided as part ofa painted aerospace component (e.g., a fuselage, an engine nacelle, awing, an empennage, a fairing and the like) having a pair of adjacentstructural panels defining therebetween a joint space having alengthwise extent and a finished paint topcoat covering an exteriorsurface of the adjacent structural panels and the joint spacetherebetween. The anti-paint cracking system will thus be associatedwith the joint space along the lengthwise extent thereof and positionedbelow the finished paint topcoat. The anti-paint cracking system willinclude a joint sealant positioned in the joint space and a polymericadhesive tape or fibrous tape formed of a fiber composite materialapplied over the joint space and the joint sealant positioned therein.

According to some embodiments, the fiber composite material willcomprise synthetic fibers embedded in a polymeric adhesive matrix. Thesynthetic fibers may be selected from the glass fibers, carbon fibers,boron fibers, and/or aramid fibers. The fiber composite materialaccording to certain specific embodiments will include glass fibersembedded in a curable polymeric resin matrix.

A primer undercoat layer may be applied to the adjacent panels under thefibrous tape so as to cover the joint sealant. Similarly, a secondaryprimer overcoat may be applied onto at least the fibrous tape below thefinished paint topcoat.

A splice plate may be rigidly connected to interior surface regions ofthe adjacent panels via suitable fasteners (e.g., rivets) along thelengthwise extent of the joint space therebetween.

These and other aspects and advantages of the present invention willbecome more clear after careful consideration is given to the followingdetailed description of the preferred exemplary embodiments thereof.

BRIEF DESCRIPTION OF ACCOMPANYING DRAWINGS

The disclosed embodiments of the present invention will be better andmore completely understood by referring to the following detaileddescription of exemplary non-limiting illustrative embodiments inconjunction with the drawings of which:

FIGS. 1A and 1B respectively depict in lesser and greater magnificationa section of a conventional painted aircraft fuselage showing cracksformed between structural panels thereof;

FIG. 2 is an exemplary schematic cross-sectional view of an anti-cracksystem in accordance with an embodiment of the invention;

FIG. 3 is a plan view of the anti-crack system shown schematically inFIG. 2 prior to application of the primer overcoat and the finishedtopcoat; and

FIG. 4 is a schematic flow diagram of a manufacturing process that maybe used to obtain the anti-crack system shown in FIGS. 2 and 3.

DETAILED DESCRIPTION

Accompanying FIGS. 1A and 1B depict the problems discussed previouslywith respect to cracks forming at the joints between structural panelsof an aerospace component (e.g., an aircraft fuselage). Specifically, asdepicted in FIGS. 1A and 1B, cracks C1 and C2 in the finished paintcoating aligned along the joints between the structural panels P1, P2and P3 are visibly perceptible. These visibly perceptible cracks C1 andC2 along with other similar cracks (not shown) in the finished paintcoating will undesirably detract from the appearance of the aircraft.

An anti-crack system 10 in accordance with an embodiment disclosedherein is schematically shown in FIG. 2. In this regard, the anti-cracksystem 10 is depicted as being applied over the joint sealant 12 whichfills the joint space between the opposed ends of structural panels 14a, 14 b, respectively. A splice plate 16 is rigidly joined to theinterior (back) side of the panels 14 a, 14 b along the lengthwisedirection (arrow L in FIG. 3) of the joint space therebetween by aseries of fasteners 18 (not shown in FIG. 2 but a representative few ofwhich are depicted in FIG. 3). Once the joint sealant 12 has beenprovided within the joint space between the panels 14 a, 14 b, a primercoating layer 20 may be applied over substantially the entire or atleast some portion of the outside (visible) surfaces laterally of thejoint space of the panels 14 a, 14 b and to cover the joint sealant 12.

Important to the embodiments disclosed herein, a strip of a fibrous tape22 is applied over and along the lengthwise direction L of the jointspace between the panels 14 a, 14 b so as to cover the initial primerundercoat layer 20. The fibrous mat 22 may be formed of virtually anysynthetic fiber material. In preferred embodiments, the fibrous tape 22is in the form of a sheet of a fiber composite material.

As used herein and in the accompanying claims, the term “fiber compositematerials” are materials that include reinforcing synthetic fibersembedded in a polymeric adhesive matrix. Fiber-reinforced compositematerials are usually supplied as fibrous sheets pre-impregnated with acurable or partially cured resin. The so-called “prepreg sheets” maythen be applied onto a structural component and cured to form rigidsheets. Fibers are embedded in a polymeric adhesive matrix (e.g., acurable epoxy resin) that connects the fibers to adjacent metal layers.Fibers may be made of, but are not limited to, glass, carbon, boron,aramid and the like. The adhesive matrix of the prepreg fiber layers maybe made of, but are not limited to, epoxy or other adhesive polymers.Preferred for use in the embodiments disclosed herein are fibercomposite materials formed of glass fibers embedded in a polymericmatrix. Especially useful are woven glass fiber tapes having a thicknessof about 0.1 mm impregnated or embedded within a curable primer epoxymatrix. The individual fibers within the fibrous tape 22 may be orientedat a bias, e.g., an angle of about 45° (+/−5-10°), relative to thelengthwise direction L of the joint space between the panels 14 a, 14 b.

The fibrous tape 22 is placed into contact on the initial primerundercoat 20 and is of sufficient width transverse to the lengthwisedirection L of the joint space between the panels 14 a, 14 b so as tooverlap opposing exterior edge regions of each such panels 14 a, 14 b.Although a single layer of fibrous tape 22 is schematically depicted inthe Figures, it is understood that multiple layers of fibrous tape 22may be applied one on top of the other as may be required. Once thefibrous tape 22 has cured in position along the lengthwise direction Lof the joint space between the panels 14 a, 14 b, a secondary primerovercoat 24 may be applied so as to cover the widthwise dimension of thetape 22. The primer overcoat 24 may thus extend beyond the lateral edgesof the tape 22 so as to contact the primer undercoat 20.

Once the primer overcoat 24 has cured, the entire surface may beprepared for the application of the finished topcoat paint layer 26,e.g., by abrading the primer undercoats 20 and 24 to enhance thedurability of topcoat layer 26.

One exemplary procedure by which the system 10 may be fabricated isdepicted in accompanying FIG. 4. As shown, surface of the panels 14 a,14 b (typically formed of an aerospace grade of aluminum but could alsobe formed of a cured fiber-reinforced composite material) may beprepared, e.g., by suitable abrasion, chemical etching and rinsing instep 30. Once the bare structural surface of the panels 14 a, 14 b hasbeen prepared, the initial primer coat 20 is applied in step 32. Thefibrous tape 34 is then applied in the lengthwise direction L of thejoint space between the panels 14 a, 14 b as previously described.Usually a single layer of fibrous tape 34 may be sufficient. Howeveralternative materials, such as auto-adhesive polyvinyl tape may beapplied according to step 36. The one or more layers of fibrous tape 34are allowed to cure in place (e.g., via ambient temperature air curing,UV light curing or heat curing depending on the particular type ofpolymeric matrix material in which the fibers are embedded). Once thefibrous tape 24 has cured, the surface may thereafter be prepared instep 38 for the topcoat finish application which is then applied in step40.

Various modifications within the skill of those in the art may beenvisioned. Therefore, while the invention has been described inconnection with what is presently considered to be the most practicaland preferred embodiment, it is to be understood that the invention isnot to be limited to the disclosed embodiment, but on the contrary, isintended to cover various modifications and equivalent arrangementsincluded within the spirit and scope thereof.

What is claimed is:
 1. A painted aerospace component comprising: a pairof adjacent structural panels defining therebetween a joint space havinga lengthwise extent; a finished paint topcoat covering an exteriorsurface of the adjacent structural panels and the joint spacetherebetween; and an anti-paint cracking system associated with thejoint space along the lengthwise extent thereof and positioned below thefinished paint topcoat, wherein the anti-paint cracking systemcomprises: (i) a joint sealant positioned in the joint space, and (ii)an adhesive tape or a fibrous tape formed of a fiber composite materialapplied over the joint space and the joint sealant positioned therein.2. The painted aerospace component as in claim 1, wherein the anti-paintcracking system comprises a fibrous tape formed of a fiber compositematerial which comprises synthetic fibers embedded in a polymericmatrix.
 3. The painted aerospace component as in claim 2, wherein thesynthetic fibers are selected from the group consisting of glass fibers,carbon fibers, boron fibers, aramid fibers and mixtures thereof.
 4. Thepainted aerospace component as in claim 3, wherein the polymeric matrixcomprises an epoxy resin.
 5. The painted aerospace component as in claim2, wherein the synthetic fibers in the fibrous tape are oriented at anangle of about 45° relative to the lengthwise extent of the joint spacebetween the adjacent panels.
 6. The painted aerospace component as inclaim 2, further comprising a primer undercoat layer applied to theadjacent panels under the fibrous tape so as to cover the joint sealant.7. The painted aerospace component as in claim 2, further comprising asecondary primer overcoat applied onto at least the fibrous tape belowthe finished paint topcoat.
 8. The painted aerospace component as inclaim 1, wherein the component is an aircraft component selected fromthe group consisting of a fuselage, an engine nacelle, a wing, anempennage and a fairing.
 9. The painted aerospace component as in claim1, further comprising a splice plate connected to interior surfaceregions of the adjacent panels along the lengthwise extent of the jointspace therebetween.
 10. An aircraft which comprises the paintedaerospace component as in claim
 1. 11. An aircraft which comprises anexterior surface painted component comprising: a pair of adjacentstructural panels defining therebetween a joint space having alengthwise extent; a splice plate connected to interior surface regionsof the adjacent panels along the lengthwise extent of the joint spacetherebetween; a finished paint topcoat covering an exterior surface ofthe adjacent structural panels and the joint space therebetween; aprimer undercoat layer applied to the adjacent panels so as to cover thejoint space along the lengthwise extent thereof; and an anti-paintcracking system associated with the joint space along the lengthwiseextent thereof and positioned below the finished paint topcoat, whereinthe anti-paint cracking system comprises: (i) a joint sealant positionedin the joint space, (ii) an adhesive tape or a fibrous tape formed of afiber composite material applied over the joint space and the jointsealant positioned therein; and (iii) a secondary primer overcoatapplied onto at least the fibrous tape below the finished paint topcoat.12. The painted aerospace component as in claim 11, wherein theanti-paint cracking system comprises a fibrous tape formed of a fibercomposite material which comprises synthetic fibers embedded in apolymeric matrix.
 13. The painted aerospace component as in claim 12,wherein the synthetic fibers are selected from the group consisting ofglass fibers, carbon fibers, boron fibers, aramid fibers and mixturesthereof.
 14. The painted aerospace component as in claim 13, wherein thepolymeric adhesive matrix comprises an epoxy resin.
 15. The paintedaerospace component as in claim 12, wherein the synthetic fibers in thefibrous tape are oriented at an angle of about 45° relative to thelengthwise extent of the joint space between the adjacent panels.
 16. Amethod to prevent crack formation along a lengthwise extent of a jointspace between painted adjacent structural panels of an aerospacecomponent, wherein the method comprises: (a) applying an adhesive tapeor a fibrous tape comprised of synthetic fibers embedded in a curablepolymeric matrix over the joint space between the adjacent structuralpanels; (b) allowing the polymeric matrix to cure; and thereafter (c)applying a paint topcoat layer over the fibrous tape and the exteriorsurfaces of the adjacent structural panels.
 17. The method as in claim16, wherein step (a) is practiced by orienting the synthetic fibers inthe fibrous tape at an angle of about 45° relative to the lengthwiseextent of the joint space.
 18. The method as in claim 16, wherein step(a) comprises inserting a joint sealant within the joint space.
 19. Themethod as in claim 16, wherein step (a) comprises applying a primerundercoat layer to the adjacent panels over the joint space, andthereafter applying the fibrous tape onto the primer undercoat layerover the joint space.
 20. The method according to claim 19, whichfurther comprises between steps (b) and (c) the step of applying asecondary primer overcoat onto at least the fibrous tape.
 21. The methodas in claim 16, wherein the fiber composite material comprises syntheticfibers embedded in a polymeric adhesive matrix.
 22. The method as inclaim 21, wherein the synthetic fibers are selected from the groupconsisting of glass fibers, carbon fibers, boron fibers, aramid fibersand mixtures thereof.
 23. The method component as in claim 22, whereinthe polymeric adhesive matrix comprises an epoxy resin.
 24. The methodas in claim 16, wherein step (a) includes orienting the synthetic fibersin the fibrous tape at an angle of about 45° relative to the lengthwiseextent of the joint space between the adjacent panels.